This invention relates to gas turbine engines and working cycles of the type in which fuel and excess air are burnt together in a first combustor, the combustion products are passed through a high pressure turbine, the exhaust of the high pressure turbine is then burnt together with further fuel in a second combustor to consume the excess air, and the exhaust of the second combustor is passed through a low pressure turbine. In particular, the invention relates to improved use of cooling air which has passed through turbine components in such a turbine.
Re-fired gas turbine engines commonly find use as prime movers in electrical generation plants. Referring to FIG. 1, in general terms an example of such an engine 1 comprises a multi-stage bladed rotary compressor 10 which compresses atmospheric air 12 to a high pressure. This compressed air 14 is then fed to a first combustor 20, which also receives gaseous and/or liquid fuel 16, the air and fuel being burnt together in a first stage of combustion. The resulting high-pressure, high-temperature combustion products 22 are used to drive a bladed rotary high pressure turbine 30, whose work output primarily drives the compressor 10 via transmission shaft 40. The air 14 supplied to combustor 20 is more than is required for complete combustion of the fuel 16 and therefore the turbine exhaust 34 contains excess air which can then be burnt with further fuel 36 in a reheat combustor 50. Bladed rotary low pressure turbine 60 receives the reheat combustion products 52 from reheat combustor 50 and uses them to drive electrical generator 70 through shaft 80. The low pressure and high pressure turbines are optionally connected together by a shaft 90, shown in dashed lines, shaft 90 being present if it is desired that both turbines 30 and 60 always run at the same speed. The exhaust 62 of the low pressure turbine 60 can be passed to atmosphere, preferably after passing through a heat exchanger for heat recovery (not shown).
To extend the life of components in the turbines, such as rotor blades, stator blades and nozzle guide vanes, it is well known to pass compressed air through them for cooling purposes. Hence, as shown in FIG. 1, both the high and low pressure turbines are provided with supplies of cooling air via lines 90 and 92 respectively, these being tapped off from the compressor 10. Because high pressure turbine 30 requires cooling air 90 which is at the highest available pressure, its supply is taken from at or near the output of the compressor 10, but low pressure turbine 60 can be supplied with cooling air 92 at a lower pressure, so it is taken from an earlier stage of the compressor. In both cases, the amount and pressure of compressed air taken from the compressor 10 for cooling purposes must be the minimum necessary to provide adequate cooling of the components, because extraction of working fluid from this early part of the engine""s working cycle imposes a cycle efficiency penalty which, unless it can be compensated for by the use of higher working fluid temperatures in the turbines, reduces the amount of power available from the low pressure turbine 60 on shaft 80.
In known engine arrangements of this type, after the cooling air has passed through the hollow interior of a turbine component such as a blade or vane, it is exhausted into the turbine annulus either from the outer tip or shroud of the blade or vane, or via small cooling holes in their flanks or in their leading or trailing edges, the cooling holes being provided in areas particularly exposed to high temperature combustion products. For example, a common cooling technique used in such circumstances is so-called xe2x80x9cfilm coolingxe2x80x9d, in which an array of small closely-spaced holes are provided to connect part of the exterior surface of the component to an interior passage through which the cooling air is flowing. Where the cooling air exits from the array of holes onto the component surface, a film of relatively cool air is formed next to the surface, thereby protecting the component from the full effects of the hot combustion gases. However, this practice of exhausting used cooling air into the turbine working passages further complicates the problem of optimizing cycle efficiency, because the cooling flow reduces the mean working fluid temperature in the turbine passage, so reducing turbine power output and efficiency.
It is an object of the present invention to increase turbine power output and turbine efficiency by reducing the diluting effects of cooling air exhausted from turbine components.
According to the present invention, a gas turbine engine comprises in flow series;
a compressor for compressing air to a high pressure,
a first combustor having fuel injection means for burning fuel together with high pressure air supplied from the compressor, the air supplied to the first combustor being in excess of that required for complete combustion of the fuel,
a high pressure turbine intended to be driven by combustion products from the first combustor,
a second combustor having fuel injection means for burning further fuel together with exhaust gases of the high pressure turbine, thereby to consume the excess air,
a lower pressure turbine having at least a first turbine stage comprising a ring of nozzle guide vanes and a first stage of rotor blades intended to be driven by combustion products from the second combustor, the lower pressure turbine having components of at least the lower pressure turbine which are cooled by cooling air supplied from the compressor,
means for supplying the second combustor with at least a portion of the cooling air after it has passed through the lower pressure turbine components, and
means for supplying sufficient fuel to the second combustor to burn therein with the portion of cooling air, thereby to increase the first stage turbine rotor entry temperature relative to an otherwise similar engine in which the portion of cooling air is exhausted into the lower pressure turbine after it has passed through the lower pressure turbine components.
For a given exit temperature of the combustion gases from the second or reheat combustor, the invention reduces cooling air dilution of the turbine gases relative to the prior art, i.e., the efficiency of the turbine is increased. For example: compare a prior art engine, in which some of the cooling air passing through the nozzle guide vanes (NGV""s) is exhausted to the turbine passage through film cooling holes, with an engine according to the present invention (but otherwise identical with the prior art), in which the same amount of cooling air is recycled to the second combustor instead of being exhausted to the turbine passage. Provided that enough fuel is supplied to be burnt with the cooling air in the second combustor to ensure the combustor exit temperature is at least maintained at the same value as in the prior art, the turbine rotor entry temperature will be increased relative to the prior art.
In an preferred embodiment of the invention, the gas turbine engine is provided with a heat exchanger arrangement for cooling the cooling air before it is supplied to the turbine components; thus, the heat exchanger arrangement may put the cooling air in heat exchange relationship with fuel, thereby to heat the fuel before injection of the fuel into the first and/or the second combustor.
The cooled turbine components may comprise stator blades in at least one stage of turbine blading and preferably include at least nozzle guide vanes constituting a first stage of the low pressure turbine.
The invention further provides a gas turbine engine operating cycle, in which fuel and excess air are burnt together in a first combustor, the combustion products are passed through a high pressure turbine, the exhaust of the high pressure turbine is then burnt together with further fuel in a second combustor to consume the excess air, and the exhaust of the second combustor is passed through a lower pressure turbine, the air supplied to the first combustor being in excess of that required for complete combustion of the fuel, wherein components of at least the lower pressure turbine are cooled by cooling air supplied from the compressor and the second combustor is supplied with at least some of the cooling air after it has passed through the turbine components, sufficient fuel being supplied to the second combustor to burn therein with the portion of cooling air, thereby to increase the first stage lower pressure turbine rotor entry temperature relative to an otherwise similar engine in which the portion of cooling air is exhausted into the first stage lower pressure turbine after it has passed through the lower pressure turbine components.
Further aspects of the invention will be apparent from the accompanying description and claims.